Method and apparatus for assembling gas turbine engine combustors

ABSTRACT

A method enables the operation of a gas turbine engine. The method comprises channeling airflow into a cooling passageway defined between the combustor casing and an inner liner of the combustor, wherein the inner liner is fabricated from a plurality of panels coupled together, channeling airflow into a cooling passageway defined between the combustor casing and an outer liner of the combustor; wherein the outer liner is fabricated from a plurality of panels coupled together, and channeling dilution airflow into a combustion chamber defined between the inner and outer liners, through a plurality of openings formed within at least one panel within at least one of the inner liner panels and the outer liner panels, wherein the plurality of openings are non-circular.

BACKGROUND OF THE INVENTION

This invention relates generally to combustors and, more particularly toa method and apparatus for decreasing combustor acoustics.

At least some known gas turbine engines include a compressor forcompressing air which is suitably mixed with a fuel and channeled to acombustor wherein the mixture is ignited for generating hot combustiongases. At least some known combustors include a dome assembly, acowling, and inner and outer liners to channel the combustion gases to aturbine, which extracts energy from the combustion gases for poweringthe compressor, as well as producing useful work to propel an aircraftin flight or to power a load, such as an electrical generator. Theliners are coupled to the dome assembly with the cowling, and extenddownstream from the cowling to define the combustion chamber. An outersupport is coupled radially outward from the outer liner such that anouter cooling passage is defined radially outward from the outer liner,and an inner support is coupled radially inward from the inner linersuch that an inner cooling passage is defined therebetween.

At least some known liners include a plurality of panels that areserially connected together between the upstream and aft ends of eachliner such that the panels define the combustion chamber. At least someknown panels are formed with primary airflow openings or secondaryairflow openings. Known primary airflow openings are formed with a firstdiameter that is sized to enable sufficient air to enter the combustionchamber to facilitate complete oxidation of the fuel within the chamber.Known secondary airflow openings are typically formed with a smallerdiameter than that of the primary airflow openings, and are positioneddownstream from the primary airflow openings. The secondary airflowopenings are sized to facilitate channeling airflow into the combustionchamber to facilitate diluting the combustion gases generated therein.However, the number of secondary openings that may be formed within agiven panel is usually limited by structural considerations, and assuch, the amount of dilution airflow that may be provided to thecombustion chamber may be limited.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, a method for operating a gas turbine engine is provided.The method comprises channeling airflow into a cooling passagewaydefined between the combustor casing and an inner liner of thecombustor, wherein the inner liner is fabricated from a plurality ofpanels coupled together, channeling airflow into a cooling passagewaydefined between the combustor casing and an outer liner of thecombustor; wherein the outer liner is fabricated from a plurality ofpanels coupled together, and channeling dilution airflow into acombustion chamber defined between the inner and outer liners, through aplurality of openings formed within at least one panel within at leastone of the inner liner panels and the outer liner panels, wherein theplurality of openings are non-circular.

In another aspect, a combustor for a gas turbine engine is provided. Thecombustor includes an inner liner, an outer liner, and a combustionchamber defined therebetween. The inner and outer liners each include aplurality of panels coupled together. At least one of the plurality ofinner liner panels and the plurality of outer liner panels includes aplurality of openings extending therethrough for channeling dilutionairflow into the combustion chamber. The plurality of openings arenon-circular.

In a further aspect, a gas turbine engine is provided. The gas turbineengine includes a combustor including an inner liner, an outer liner,and a combustion chamber defined therebetween. The inner and outerliners each include a plurality of panels coupled together. At least oneof the plurality of inner liner panels and the plurality of outer linerpanels includes a plurality of openings extending therethrough forchanneling dilution airflow into the combustion chamber. The pluralityof openings are non-circular.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration of an exemplary gas turbine engine;

FIG. 2 is a cross-sectional view of a combustor that may be used withthe gas turbine engine;

FIG. 3 is an enlarged perspective view of a portion of a liner used withthe combustor shown in FIG. 2 and taken along area 3; and

FIG. 4 is a plan view of a portion of the liner used with the combustorshown in FIG. 2 and taken along area 4.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 is a schematic illustration of an exemplary gas turbine engine 10including a low pressure compressor 12, a high pressure compressor 14,and a combustor 16. Engine 10 also includes a high pressure turbine 18,and a low pressure turbine 20 arranged in a serial, axial flowrelationship. Compressor 12 and turbine 20 are coupled by a first shaft24, and compressor 14 and turbine 18 are coupled by a second shaft 26.In one embodiment, gas turbine engine 10 is an LMS100 enginecommercially available from General Electric Company, Cincinnati, Ohio.

In operation, air flows through low pressure compressor 12 from anupstream side 28 of engine 10. Compressed air is supplied from lowpressure compressor 12 to high pressure compressor 14. Highly compressedair is then delivered to combustor assembly 16 where it is mixed withfuel and ignited. Combustion gases are channeled from combustor assembly16 to drive turbines 18 and 20.

FIG. 2 is a cross-sectional view of a combustor 30 that may be used withgas turbine engine 10. FIG. 3 is an enlarged perspective view of aportion of a liner 40 used with combustor 30 and taken along area 3.FIG. 4 is a plan view of a portion of liner 40 used with combustor 30shown in FIG. 2 and taken along area 4. Combustor 30 includes a domeassembly 32. A fuel injector 34 extends into dome assembly 32 andinjects atomized fuel through dome assembly 32 into a combustion zone orchamber 36 of combustor 30 to form an air-fuel mixture that is igniteddownstream of fuel injector 34.

Combustor dome assembly 32 defines an upstream end of combustion zone 36and includes a plurality of mixer assemblies 37 that are spacedcircumferentially around combustor dome assembly 32 for delivering amixture of fuel and air to combustion zone 36. In the exemplaryembodiment, combustor dome assembly 32 is a single annular combustor(SAC) that includes one annular combustor dome. However, it should beunderstood that in alternative embodiments combustor dome assembly 32may include any number of combustor domes. For example, in oneembodiment, combustor dome assembly 32 is a dual annular combustor(DAC), and, in another embodiment, combustor dome assembly 32 is atriple annular combustor.

Combustion zone 36 is defined by combustor liners 40 that shieldcomponents external to combustor 30 from heat generated withincombustion zone 36. Combustion zone 36 extends from dome assembly 32downstream to a turbine nozzle assembly 41. Liners 40 include an innerliner 42 and an outer liner 44. Each liner 42 and 44 is annular andincludes a plurality of separate panels 50. In the exemplary embodiment,each panel 50 includes a series of steps 52, each of which form adistinct portion of combustor liner 40.

Outer liner 44 and inner liner 42 each include a respective aft-mostpanel 64 and 66. Panels 64 and 66 are each located at the aft end 68 ofcombustion zone 36 and are adjacent turbine nozzle assembly 41.Specifically, each panel 64 and 66 couples an aft end 70 and 72 of eachrespective liner 44 and 42 to turbine nozzle assembly 41.

Each liner 42 and 44 also includes an annular support mount, or aftmount, 80 and 82, respectively. Specifically, each support mount 80 and82 couples an aft end 70 and 72 of each respective liner 44 and 42 toturbine nozzle assembly 41 and to a combustor casing 84 that extendssubstantially circumferentially around combustor 30. More specifically,each support mount 80 and 82 extends radially outward from eachrespective liner 42 and 44 such that a radially outer cooling passageway86 and a radially outer cooling passageway 88 are defined betweencombustor casing 84 and combustor liner 40. Accordingly, coolingpassageway 86 is defined between liner 42 and combustor casing 84 andcooling passageway 88 is defined between liner 44 and combustor casing84.

Each combustor panel 50 includes a combustor liner surface 90 and anexterior surface 92 that is radially outward from liner surface 90. Whenpanels 50 are coupled together, combustor liner surface 90 extendsgenerally from dome assembly 32 to turbine nozzle assembly 41. In theexemplary embodiment, each panel 50 is generally rectangular andincludes a pair of circumferentially-spaced side edges 100 that areconnected together by a leading edge side 102 and an opposed trailingedge side 104.

Each liner 42 and 44 also includes at least one panel 110 that isdownstream from fuel injector 34, and includes a plurality ofcircumferentially-spaced primary airflow openings 111 that extendthrough panel 110 between combustor liner surface 90 and an exteriorsurface 92. Openings 111 are substantially circular and have a diameterD₁. In the exemplary embodiment, openings 111 extend substantiallycircumferentially around combustion chamber 36. Accordingly, openings111 connect each cooling passageway 86 and 88 in flow communication withcombustion chamber 36. In the exemplary embodiment, panel 110 is atleast two panels 50 upstream from turbine nozzle assembly 41.

Each liner 42 and 44 also includes at least one panel 112 that isdownstream from panel 110 and includes a plurality ofcircumferentially-spaced secondary or dilution airflow openings 116. Inthe exemplary embodiment, openings 116 are spaced substantiallycircumferentially around combustion chamber 36. Openings 116 extendthrough panel 112 between combustor liner surface 90 and an exteriorsurface 92 and are non-circular. More specifically, in the exemplaryembodiment, openings 116 are substantially race-tracked shaped orgenerally elliptical and are defined by a pair of opposed, generallyparallel sidewalls 120 that are connected by a pair of opposed arcuatesidewalls 122.

In the exemplary embodiment, sidewalls 122 are formed with apre-determined radius of curvature R₁ that is smaller than an associatedradius 1/2 D₁ of each primary cooling opening 111. More specifically, inthe exemplary embodiment, each sidewall 122 is substantiallysemi-circular. Accordingly, because sidewalls 120 are substantiallyparallel, within each opening, sidewalls 120 are separated by thediameter D₃ (twice the radius of curvature R₁) of each arcuate sidewall122.

In the exemplary embodiment, openings 116 are oriented such thatsidewalls 120 are aligned generally axially. Accordingly, a distance ofseparation, known as web spacing, D₂ between circumferentially adjacentopenings 116 is measured between adjacent opening sidewalls 120. In theexemplary embodiment, distance D₂ is at least twice that of the diameterD₃ of each opening 116. The distance of separation D₂ facilitatesmaintaining structural integrity of each panel 112.

During operation, an annular diffuser 124 channels air discharged fromcompressor 14 into the combustor dome assemblies 32 and, morespecifically, into mixer assemblies 37, wherein the compressed air ismixed with fuel provided by fuel injector 34. The fuel/air mixture isthen ignited within combustion zone 36 to form combustion gases, whichare discharged from the combustion zone 36 through turbine nozzleassembly 41.

A portion of the compressed air discharged from compressor 14 ischanneled into each cooling passageway 86 and 88 for cooling combustorassembly 30. More specifically, the compressed air channeled throughpassageways 86 and 88 is also channeled into combustion zone 36 throughprimary cooling openings 111 defined within panels 110. The compressedair channeled through openings 111 facilitates convectively coolingliners 42 and 44 in regions adjacent openings 111. Moreover, airchanneled through openings 111 also mixes with the fuel-air mixturewithin combustion chamber 36 to facilitate complete oxidation of all ofthe fuel supplied to chamber 36.

As the fuel-air mixture is channeled downstream, the mixture is mixedwith air channeled through dilution openings 116. Openings 116facilitate diluting the burned combustion products within chamber 36 tofacilitate reducing the temperature of the combustion gases channeleddownstream to the turbines. Moreover, the elongation of openings 116facilitates increasing the penetration of the airflow jets dischargedinto chamber 36 from openings 116 in comparison to other known dilutionopenings, such as circular openings. The increased penetration of thedilution airflow enables openings 116 to facilitate shaping the radialtemperature profile to a predetermined profile shape in an area wherethe mainstream velocities are relatively high.

The elongated shape of openings 116 facilitates enough penetration ofthe dilution air such that the air is not readily turned or forced overby the mainstream flow. Moreover, because openings 116 are oriented withtheir narrowest dimension in the circumferential direction, airflowdischarged through openings 116 is more streamlined than airflowdischarged through circular openings, which enables the airflow topenetrate the mainstream flow to a greater extent than is possiblethrough a round opening.

The above-described gas turbine engine combustor includes a liner thatincludes at least one panel including a plurality of non-circular,dilution openings extending therethrough. The dilution openings areoriented such that their narrowest dimension extends circumferentiallyacross the panel. The shape and orientation of the dilution openingsenables airflow discharged from the openings to penetrate the mainstreamflow to a greater extent than is possible through known round openings.As a result, the openings facilitate enhanced control of the radialtemperature profile generated within the combustion chamber andincreasing the useful life of the combustor in a cost-effective andreliable manner.

Exemplary embodiments of a combustor for a gas turbine engine aredescribed above in detail. The systems and assembly components of thecombustor are not limited to the specific embodiments described herein,but rather, components of each system may be utilized independently andseparately from other components described herein. Each system andassembly component can also be used in combination with other combustorsystems and assemblies or with other gas turbine engine components.

While the invention has been described in terms of various specificembodiments, those skilled in the art will recognize that the inventioncan be practiced with modification within the spirit and scope of theclaims.

1. A method for operating a gas turbine engine, said method comprising:channeling airflow into a cooling passageway defined between thecombustor casing and an inner liner of the combustor, wherein the innerliner is fabricated from a plurality of panels coupled together;channeling airflow into a cooling passageway defined between thecombustor casing and an outer liner of the combustor; wherein the outerliner is fabricated from a plurality of panels coupled together; andchanneling dilution airflow into a combustion chamber defined betweenthe inner and outer liners, through a plurality of openings formedwithin at least one panel within at least one of the inner liner panelsand the outer liner panels, wherein the plurality of openings arenon-circular.
 2. A method in accordance with claim 1 wherein channelingdilution airflow into a combustion chamber further comprises channelingdilution airflow into the combustion chamber to facilitate controllingan exit temperature profile of the combustor.
 3. A method in accordancewith claim 1 wherein channeling dilution airflow into a combustionchamber further comprises channeling dilution airflow into thecombustion chamber through the plurality of openings, wherein theopenings are shaped to enable cooling air to penetrate into thecombustion chamber to facilitate achieving a desired radial temperatureprofile within the combustion chamber.
 4. A method in accordance withclaim 1 wherein channeling dilution airflow into a combustion chamberfurther comprises channeling dilution airflow into the combustionchamber through the plurality of openings, wherein the openings aregenerally elliptically shaped.
 5. A method in accordance with claim 1wherein channeling dilution airflow into a combustion chamber furthercomprises channeling dilution airflow into the combustion chamberthrough the plurality of openings, wherein the openings are defined by apair of substantially parallel walls that are connected together by apair of opposed arcuate sidewalls formed with a predetermined radius ofcurvature.
 6. A combustor for a gas turbine engine, said combustorcomprising: an inner liner comprising a plurality of panels coupledtogether; an outer liner comprising a plurality of panels coupledtogether; and a combustion chamber defined between said inner and outerliners, at least one of said plurality of inner liner panels and saidplurality of outer liner panels comprises a plurality of openingsextending therethrough for channeling dilution airflow into saidcombustion chamber, said plurality of openings are non-circular.
 7. Acombustor in accordance with claim 6 wherein said plurality of openingsfacilitate controlling an exit temperature profile of said combustor. 8.A combustor in accordance with claim 6 wherein said plurality ofopenings are each substantially elliptically-shaped.
 9. A combustor inaccordance with claim 6 wherein said plurality of openings are shaped toenable cooling air to penetrate into said combustion chamber tofacilitate achieving a desired radial temperature profile within saidcombustion chamber.
 10. A combustor in accordance with claim 6 whereinsaid plurality of openings are defined by a pair of opposedsubstantially parallel sidewalls connected together by a pair of opposedarcuate walls formed with a pre-determined radius.
 11. A combustor inaccordance with claim 10 wherein adjacent of said plurality of openingsare separated by a distance that is approximately equal to twice thediameter of said arcuate walls.
 12. A combustor in accordance with claim6 wherein said at least one panel comprises a pair of opposedcircumferential edges coupled together by a leading edge and a sideedge, said plurality of openings comprises at least three openingsspaced approximately equi-distantly between said pair of opposedcircumferential edges.
 13. A gas turbine engine comprising a combustorcomprising an inner liner, an outer liner, and a combustion chamberdefined between said inner and outer liners, each of said inner andouter liners comprises a plurality of panels coupled together, at leastone of said panels within at least one of said inner liner and saidouter liner comprises a plurality of openings extending therethrough forchanneling dilution air into said combustion chamber, said plurality ofopenings are non-circular.
 14. A gas turbine engine in accordance withclaim 13 wherein said combustor plurality of openings extending throughsaid at least one panel facilitate controlling an exit temperatureprofile of said combustor.
 15. A gas turbine engine in accordance withclaim 14 wherein said combustor plurality of openings extending throughsaid at least one panel are each generally elliptically-shaped.
 16. Agas turbine engine in accordance with claim 14 wherein said combustorplurality of openings extending through said at least one panel areshaped to enable cooling air to penetrate into said combustion chamberto facilitate achieving a desired radial temperature profile within saidcombustion chamber.
 17. A gas turbine engine in accordance with claim 14wherein said combustor plurality of openings extending through said atleast one panel are defined by a pair of opposed substantially parallelsidewalls that are connected together by a pair of opposed arcuate wallsformed with a pre-determined radius.
 18. A gas turbine engine inaccordance with claim 17 wherein adjacent of said plurality of openingsextending through said at least one panel are separated within saidpanel by a distance that is approximately equal to twice the diameter ofsaid arcuate walls.
 19. A gas turbine engine in accordance with claim 14wherein each of said plurality of panels comprises a pair of opposedcircumferential edges coupled together by a leading edge and a sideedge, said plurality of openings extending through said at least onepanel comprises at least three openings spaced approximatelyequi-distantly between said pair of opposed circumferential edges.